The present invention relates generally to heat conducting devices, and, more particularly, to passive heat conduction devices that act analogously to electrical diodes, that is, allowing heat to pass in a first configuration and preventing heat from passing in a second configuration.
Circuit boards in missiles contain electronic components mounted on and through the circuit board. The electronics themselves generate heat, and the circuit board may also experience an influx of aerodynamic heat through the airframe of the missile during high speed flight. Both internal and external sources of heat may degrade the operation of the circuit board and the electronics, by increasing their operating temperature.
In many older missiles, the influx of heat during high speed flight was the primary concern, because the electronics did not produce large amounts of heat during limited flight times. Accordingly, missiles that had short-duration, high-speed flights were designed to limit the influx of heat during that mode of operation. These designs, however, did not allow for the removal of heat from the missile. Thus, as the electronics have become more complex and have begun to operate for longer periods of time, internal heat build-up has become a concern, as well. Thus, circuit boards in some modern missiles may generate significant internal heat and also experience significant heating from the airframe during high speed flight.
Specifically, present-day missile systems are being asked to operate under harsh environments with increasingly high power densities and longer operating times. For reliability, electronic components must be maintained below certain temperature limits (typically 110xc2x0 Celsius junction temperatures). Increasing power density and increasing performance requirements make staying within these limits increasingly difficult. Many of the environments under which a missile is asked to operate for extended periods provide the opportunity to reject this heat through the airframe, or missile skin. Such a heat rejection requires a good thermal path to the missile skin. Examples include production testing, high altitude/low speed captive carry, and field reprogramming. The ultimate operational environment is, however, very harsh thermally, providing high heating to the missile skin. Examples include high speed dash in captive carry of airborne missiles and free flight in most all missiles. This environment calls for thermal isolation of the internal components from the missile skin.
The goals of a good thermal path to the missile skin for long-term operation and thermal isolation for short term operation in free flight are directly at odds, unless a thermal diode can be developed which allows unrestricted heat-flow in one direction (out of missile) and highly restricted heat flow in the other direction (into missile).
Application Ser. No. 09/389,655, filed on Sep. 2, 1999, entitled xe2x80x9cHeat Conducting Device for a Circuit Boardxe2x80x9d and assigned to the same assignee as the present invention, discloses and claims one solution to the above-discussed problems. That solution comprises a heat conducting device with a low thermal impedance to the environmental sink for the heat generated by the electronics in the circuit board in one mode of operation of the missile and a high thermal impedance to environmental source for the heat generated by the aerodynamics on the airframe in another mode of operation of the missile. Thus, the claimed solution comprises a thermal diode. However, while the claimed invention is certainly useful for its intended purpose, work continues in an effort to develop improved thermal diodes. In particular, because the ""655 application is not a passive device and requires some sort of control logic, it is desirable to provide a passive, self-contained sensor-actuator system that does not require a separate sensor, control system, and actuator.
In accordance with the present invention, a passively operated thermal diode for controlling heat transfer from heat-generating electronic components to an external environment through an airframe, or missile skin, is provided. The thermal diode comprises:
an electronics package within the airframe;
a heat-transferring mechanism thermally connected to the electronics package and controllably disengagable from thermal contact with the airframe at a predetermined temperature;
a shape-memory alloy component having a phase change at the predetermined temperature for thermally disconnecting the electronics package from the airframe once the predetermined temperature is exceeded; and
a spring for maintaining the heat-transferring mechanism in thermal contact the airframe below the predetermined temperature.
In one embodiment, the electronics package is provided with at least one first tapered surface. The heat-transferring mechanism is provided with at least one second tapered surface that slidably mates with the first tapered surface(s) and a cylindrical face that mates with the missile skin. The shape-memory component comprises an actuator for moving the heat-transferring mechanism out of thermal contact with the missile skin once the predetermined temperature is exceeded. The spring, e.g., a Belleville spring, maintains the heat-transferring mechanism in thermal contact with the first tapered surface and the missile skin below the predetermined temperature.
In a second embodiment, a portion of an inner surface of the airframe is provided with two tapered surfaces, a first tapered surface having a comparatively large contact area and a comparatively shallow taper and a second tapered surface having a comparatively small contact area and a comparatively steep taper. The electronics package is thermally connected to a heat sink also provided with two tapered surfaces, a third tapered surface having a comparatively large contact area and a comparatively shallow taper and a fourth tapered surface having a comparatively small contact area and a comparatively steep taper. The heat sink is configured (1) with its third tapered surface opposed to the first tapered surface and with its fourth tapered surface opposed to the second tapered surface and (2) to slidably move from contact of the first and third surfaces to contact of the second and fourth surfaces. The shape-memory alloy component comprises a washer for moving the heat sink out of thermal contact with the first tapered surface once the predetermined temperature is exceeded. The spring comprises a spring-loaded washer for maintaining the heat sink in thermal contact with the first tapered surface below the predetermined temperature.
The present invention solves the problem of heat transfer by introducing a shape-memory-alloy sensor-actuator component. The SMA component deforms due to a temperature-initiated phase change, thereby providing the force and motion required to change the mechanical connection between the missile avionics and the airframe from a thermally conductive path to a thermally insulating path.
Thus, the present invention provides a passive, self-contained sensor actuator system that does not require a separate sensor, control system, and actuator.
The application of this technology to the thermal management of missile electronics has significant potential for all missiles. Thermal management of missile electronics must strike the balance between good coupling with the environment (for captive carry, flight line testing, and production test) and isolation from the environment (to minimize the impact of aerothermal heating during free flight). Traditionally, this has been accomplished by tailoring thermal conductance paths and managing thermal mass to passively achieve the goals of survival during long-term/low-temperature and short-term/high-temperature operating conditions. However, as the power density of electronics packages increases, this approach is not adequate, and more advanced techniques must be developed.
The approach disclosed herein offers a solution that has the potential to greatly expand the allowable operating envelopes of missile electronics. In terms of missile performance, this means;
Longer allowable run-times and shorter cooldown times during operation, leading to decreased cycle time.
Higher allowable electronics power dissipation associated with increased functionality.
Expanded allowable captive-carry flight speed and endurance envelope.
Increased allowable flight time and speed which increase stand-off distance.
The present invention is a passive extension and improvement to the invention disclosed and claimed in the above-referenced related patent application. However, it is very different from previously implemented missile electronics packaging techniques in that most previous techniques use isolation of the internal components in order to survive free flight. This, however, limits captive carry and ground test operation, and requires long non-operational cool down times. Rather, the present invention extends the system of the related patent application because it is passive and does not require an active sensor, control and actuation system.